1. Field of the Invention
The present invention relates to systems for sensing the motion of a vehicle. Particularly, the present invention relates to systems and techniques for sensing satellite nutation.
While the present invention is described herein with reference to a particular embodiment for a particular application, it is understood that the invention is not limited thereto. Those of ordinary skill in the art will recognize additional embodiments and applications within the scope thereof.
2. Description of the Related Art
The advantages of spin stabilization for satellite attitude control have been repeatedly demonstrated. Unfortunately, spin stabilized satellites are inherently susceptible to nutation. That is, in the absence of external torques, the angular momentum vector of a spin stabilized satellite would be fixed in inertial space. Nutation, a coning or precessing motion of the spin axis, fixed in the body, around the angular momentum vector, results from the misalignment of the spin axis by a transverse angular momentum. Transverse angular momentum may be induced by the firing of control thrusters during attitude and orbit correction maneuvers; by the motion of articulated payload elements; by the effects of flexible elements; or by the sloshing of liquids in the spacecraft.
Whatever the cause, accurate spin stabilized satellite attitude control requires the damping of nutation. To this end, nutation sensors are used in conjunction with mechanical elements located on the spacecraft. The mechanical elements provide the energy dissipation (or addition) required to reduce the wobble by removing the transverse angular momentum vector.
Nutation sensors are typically linear accelerometers mounted on the rim of the spacecraft. The accelerometer is often a hinged pendulus mass, mounted to sense the up and down motion, due to nutation, along an axis parallel to the spin axis.
Current nutation sensors have limited low frequency response. As a result, such sensors have difficulty measuring very slow nutation frequencies as may be experienced by large structures having low spin rates. For such applications, it is desirable to provide an accurate low speed nutation sensor.
A second shortcoming of current nutation sensors is that they tend to misinterpret the acceleration due to the firing of the thrusters as nutation. Accordingly, the output of the sensor is invalid, if uncorrected, during thruster firings. This is undesirable as information due to nutation during thruster firing may be used to provide for greater control and efficiency during such manuevers. This in turn would allow for lower fuel costs and less time to execute a particular maneuver. Thus it is generally desirable to provide a nutation sensor which is accurate during thruster firings.
A third shortcoming of many current nutation sensors is that for large spacecraft having low spin rates, the accelerometer sensor must be placed far from the spin axis to develop a sufficiently strong output signal. An example of such an spacecraft is the Space Station currently under development by NASA.
As the distance from the spin axis increases, however, the effects of flexibility have a more significant impact on the sensor output. Thus, under some circumstances, the flexure of the spacecraft may be interpreted as nutation. It is generally desirable therefore to provide a nutation sensing system which provides an output independent of the flexure modes of the spacecraft.